This invention relates generally to gas turbine engines, and more particularly to apparatus and methods for actively controlling the radial clearances between rotors and shrouds in the turbine sections of such engines.
A typical gas turbine engine includes a turbomachinery core having a high pressure compressor, a combustor, and a high pressure turbine in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. The high pressure turbine or (“HPT”) includes one or more rotors which extract energy from the primary gas flow. Each rotor comprises an annular array of blades or buckets carried by a rotating disk. The flowpath through the rotor is defined in part by a shroud, which is a stationary structure carried by a turbine case and which circumscribes the tips of the blades or buckets. These components operate in an extremely high temperature environment.
Blade tip clearances are a critical component of overall engine performance, especially the tip clearances in the HPT. Because gas turbine engines operate over a wide range of operating conditions, it is generally not possible to set the static blade tip clearances so as to maintain best efficiency while also avoiding “rubs” between the blade tips and the surrounding structure at all engine operating conditions. It is therefore known to actively control blade tip clearance by selectively heating and/or cooling the turbine case.
However, such systems are typically dependent on the use of complex, expensive manifold structures to deliver the heating or cooling air to the turbine case, and also require complex valving and piping to control the extraction and delivery of high-pressure bleed air to the manifolds.
Accordingly, there is a need for a means of providing active clearance control in a gas turbine engine with minimum weight and expense.